Engine control system

ABSTRACT

A power turbine speed control system for a helicopter is disclosed which includes components for generating a power turbine speed signal based upon a demanded rotor speed, a high-order filter for filtering the power turbine speed signal by effectuating a rapid attenuation of main and tail rotor torsional frequencies in the power turbine speed signal without compromising phase at low frequencies, and a governor for providing isochronous power turbine speed and rotor speed control based upon the filtered power turbine speed signal.

CROSS-REFERENCE TO RELATED APPLICATION

The subject application is a divisional application of U.S. patentapplication Ser. No. 09/963,180 filed Sep. 26, 2001 now U.S. Pat No.6,729,139.

GOVERNMENT RIGHTS STATEMENT

The U.S. Government has a paid-up license in this invention and theright in limited circumstances to require the patent owner to licenseothers on reasonable terms as provided for by the terms ofDAAH10-99-2-0005, awarded by the U.S. Department of the Army.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The subject invention relates to the operation of gas turbine engines,and more particularly, to a control system for a turboshaft engine of ahelicopter which includes a wide-band speed governor.

2. Background of the Related Art

In rotary wing aircraft, and more particularly, in helicopters, the mainand tail rotors define the primary flight control surfaces for theaircraft. The rotor drive train for the main and tail rotor is coupledto the power plant which can have a single or twin engine configuration.Engine response, is therefore critical to the control of the aircraft.Even more critical to the control of the aircraft is fast fuel flowresponse to the engine. Thus, it is desirable to provide a rotor speedcontrol system with a high bandwidth enabled by increased proportionaland derivative power turbine governing gains. This can be achieved byappropriate filtering of the speed feedback signal.

It is also desirable to provide a rotor speed control system that hasincreased main and tail rotor resonant frequency attenuation as comparedto prior art control systems. It would also be desirable to provide arotor speed control system that eliminates the need for yaw and lateralcyclic load anticpation.

These and other desirable attributes are achieved as part of the subjectinvention by providing a high-order filter in the speed control loop ofthe engine control system disclosed herein.

SUMMARY OF THE INVENTION

The subject invention is directed to a new and useful control system fora turboshaft engine of a helicopter. The control system includes a speedcontrol loop with means for generating a power turbine speed signalbased upon a demanded rotor speed, means for filtering the power turbinespeed signal by effectuating a rapid attenuation of main and tail rotortorsional frequencies in the power turbine speed signal withoutcompromising phase at low frequencies, and a governor for providingisochronous power turbine speed and rotor speed control based upon thefiltered power turbine speed signal.

In accordance with the subject invention, the means for filtering thepower turbine speed signal is a high-order filter. Preferably, thehigh-order filter is an eighth order filter. In an embodiment of theinvention, the high-order filter is configured as three second orderfilters cascaded in series with two first order filters. Alternatively,the high-order filter is configured as a sixth order filter in serieswith a second order filter.

The speed control system of the subject invention further includes anoptional active torsional damping loop for damping main and tail rotortorsional frequencies. The torsional damping loop includes a Kalmanstate estimator for estimating a plurality of engine states based uponone or more measured engine states, such as, for example power turbineshaft torque (QS). The torsional damping loop also includes a LinearQuadratic Regulator (LQR) that provides combustive damping.

These and other aspects of the subject invention will become morereadily apparent to those having ordinary skill in the art from thefollowing detailed description of the invention taken in conjunctionwith the drawings described hereinbelow.

BRIEF DESCRIPTION OF THE DRAWINGS

So that those having ordinary skill in the art to which the subjectinvention pertains will more readily understand how to employ the enginecontrol system of the subject invention, preferred embodiments thereofwill be described in detail hereinbelow with reference to the drawings,wherein:

FIG. 1 is a schematic representation of the control system of thesubject invention including an outer speed control loop with ahigh-order torsional filter, and an optional inner torsional dampingloop;

FIG. 2 is a bode diagram comparing a baseline notch/lag filter to thehigh-order filter of the subject invention;

FIG. 3 is a baseline control open loop bode diagram using a notchfilter;

FIG. 4 is a wide-band governor open loop bode plot using a high-orderattenuating filter; and

FIG. 5 is a diagram illustrating the signal flow through the optionaltorsional damping loop of the control system of FIG. 1.

DETAILED DESCRIPTION OF PREFFERED EMBODIMENTS

Several engine operating parameters are associated with a turboshaftengine for rotary wing aircraft and will be referred to in thedescription of the invention. These operating parameters include thefollowing:

NF* Demanded Free Power Turbine Speed NF Free Power Turbine Speed NG GasGenerator Speed NDOT* Demanded Rate of Change of Power Turbine Speed WF*Demanded Fuel Flow WF Actual Fuel Flow P3 Compressor Discharge PressureHMU Hydro-mechanical Unit QGAS Gas Generator Output Torque QS PowerTurbine Shaft Torque

Referring now to the drawings wherein like reference numerals identifysimilar features of the invention disclosed herein, there is illustratedin FIG. 1 a schematic representation of the engine control system of thesubject invention which is designated generally by reference numeral 10.More particularly, the engine control system 10 includes an outer enginespeed control loop 12 and an inner torsional damping loop 14 that may beselectively activated by the pilot of the aircraft. In accordance withthe subject invention, the inner torsional damping loop 14 providesactive damping of torsional resonance frequencies without significantlyaffecting the rotor speed governing loop 12 of the engine control system10.

Speed Control Loop

The outer engine speed control loop 12 includes a Power Turbine Governor(PTG) 16 providing proportional and derivative control paths. The PTGreceives an input signal from a primary summing junction 18. This signalis based upon the pilot's demanded rotor speed (NF*) and the filteredfree turbine speed signal (NF). The PTG delivers an output signal to asecondary summing junction 20 which also receives input signals fromrotor load anticipators and the optional torsional damping loop 14 whenactivated by the pilot. The resultant signal from summing junction 20 isindicative of the Demanded Rate of Change of the Gas Generator Speed(NDOT*). This signal is input into an auctioneering circuit 22 whichincorporates software programming that determines a highest or lowestNDOT* value based upon engine limits related toacceleration/deceleration, torque and temperature.

A signal indicative of the winning NDOT* value is used as the input tothe fuel control unit 24, together with a signal indicative of the gasgenerator speed (NG) sensed from the core engine 30. Based upon thesesignals, the fuel control unit 24 generates a proportional signalindicative of the fuel flow (WF) divided by the compressor dischargepressure (P3). This proportional signal is input into a multiplierjunction 26 that receives a sensed signal indicative of the compressordischarge pressure (P3) of the core engine 30. As a result, the (P3)denominator is dropped from the proportional signal by the multiplier,and the resulting output signal from multiplier junction 26 isindicative of the demanded fuel flow (WF*).

A signal indicative of the demanded fuel flow (WF*) is input into thehydro-mechanical unit (HMU) 28 that includes, among other things, a fuelmetering valve (not shown) for regulating the amount of fuel deliveredto the core engine 30. Thus the demanded fuel flow corresponds to ademanded metering valve position.

A signal indicative of the engine output torque (QGAS) is transmittedfrom the core engine 30 to the power turbines and rotor drive traincontrol unit 32, which, in turn produces two distinct signals. Onesignal is indicative of the free power turbine speed (NF), and the othersignal is indicative of the power turbine shaft torque (QS). The signalindicative of NF is input into a high-order torsional filter 40, thefunction and configuration of which is described in detail hereinbelow.The signal indicative of QS is used as the input signal to the optionaltorsional damping loop 14, as will also be discussed hereinbelow withreference to FIG. 5.

Those skilled in the art will readily appreciate that the signalindicative of NF is a compound signal containing a plurality ofcontributing frequencies. Included among these frequencies, are the mainand tail rotor torsional frequencies. The high-order filter 40 isadapted and configured to attenuate torsional oscillations in the powerturbine speed signal (NF) that are directly associated with the main andtail rotors. Moreover, as illustrated in FIG. 2 and explained in Table1, the high-order filter 40 provides greater attenuation of the main andtail rotor resonant frequencies, as compared to a baseline controlsystem, in both single engine and twin engine configurations. This isexhibited by the sharp gain roll-off at the main and tail rotor resonantfrequencies depicted in the upper bode plot of FIG. 2. As shown, thehigh-order filter also allows for the magnitude peak at about 30rad/sec. Furthermore, as illustrated in the lower bode plot of FIG. 2,the high-order filter 40 functions to minimize the low frequency phaselag, as compared to a baseline control system. The increased attenuationprovided by the high-order filter allows the proportional and derivativegains of the PTG to be increased by about a factor of three, as comparedto a baseline control system. Consequently, the bandwidth of the outerturbine speed control loop 12 is increased by about a factor three ascompared to the baseline control system.

Table 1.0 summarizes the differences in characteristics of the baselineand high-order torsional filter for single and twin engines coupled tothe UH-60L Black Hawk rotor drive train.

TABLE 1.0 Main Rotor Resonance Tail Rotor Resonance High-Order BaselineHigh-Order Baseline Filter Notch/Lag Filter Notch Lag FrequencyAttenuation Attenuation Frequency Attenuation Attenuation One Engine17.75 rad/s 32.2 dB 21.2 dB 42.50 rad/s 44.5 dB 26.1 dB Twin Engine18.85 rad/s 28.0 dB 20.5 dB 37.00 rad/s 40.7 dB 24.9 dB

Referring now to FIGS. 3 and 4, there are illustrated open loop bodediagrams for twin engine operation at sea level, standard day ambientconditions for the baseline and wide-band governor system. Asillustrated in the upper plot of FIG. 3, the cross-over frequency atunity gain or the bandwidth of the baseline control system is about 3–4rad/sec. In contrast, as illustrated in the upper plot of FIG. 4, thecross-over frequency or bandwidth of the wide-band governor is about8–10 rad/sec. This represents an increase in bandwidth on the order oftwo to three times that of the baseline system.

The bode diagrams at altitudes of 4000 and 10,000 feet have similarshapes with characteristics shown in the following tables. Table 2.1presents a summary of the baseline system's power turbine governorcontrol loop stability margins and open loop system bandwidth for twinengine operation when coupled to the rotor drive train. The tail rotortorsional attenuation exceeds that of the main rotor, therefore thetorsional attenuation listed in the table represents that of the mainrotor.

TABLE 2.1 Gain Margin Phase Margin Open loop Main Rotor Alt. (dB) (°)Bandwidth Attenuation (ft) (° F.) Min Max Min Max Min Max Min Max 0 595.8 6.8 56.9 91.2 3.07 4.18 5.3 6.4 4K 95 6.3 7.0 53.5 87.2 3.01 4.146.0 6.5 10K 23.3 6.6 8.0 67.3 95.2 2.05 3.06 6.2 7.2

Table 2.2 presents a summary of the wide-band governor control loopstability margins and open loop system bandwidth for twin engineoperation when coupled to the rotor drive train. The increased open loopbandwidth is due to tripling the PD gains, It should be noted hat thereis no significant decrease in the main torsional frequency attenuationdespite the increase in bandwidth.

TABLE 2.2 Gain Margin Phase Margin Open loop Main Rotor Alt. (dB) (°)Bandwidth Attenuation (ft.) (° F.) Min Max Min Max Min Max Min Max 0 598.4 9.6 53.7 79.4 5.45 11.52 5.5 6.6 4K 95 9.2 13.0 50.9 81.3 5.42 10.325.7 6.9 10K 23.3 10.3 14.6 57.0 82.3 5.06 5.60 5.7 7.8

The analysis was performed using a simulated Comanche-style steppermotor based fuel metering system as well as a variable displacement vanepump (VDVP) based fuel metering system operating with twin PWC 3000 SHP2.5 second engines as well as twin empirical 3000 SHP one-secondengines.

The wide-band governor of the subject invention runs on an ECU with asampling time of ten milliseconds. The high-order filter allows for atripling of proportional and derivative (PD) gains when compared to thebaseline control. The wide-band governor of the subject invention isconfigured to maintain open loop gain and phase margins of at least 6 dBand 45° respectively, throughout the powered region of the engineoperating envelope when coupled to the rotor drive train. Thisconfiguration provides stable, responsive and well damped power turbinespeed control.

The baseline speed control loop incorporates, in the feedback path, asecond order attenuating lead-lag (notch) filter, sized for the BlackHawk helicopter, cascaded in series with a first order lag filter toprovide gain attenuation of the main and tail rotor resonant modes.

In contrast, the wide-band governor speed control loop of the subjectinvention incorporates, in the feedback path, an eighth orderattenuating filter to provide increased gain attenuation of the main andtail rotor resonant modes. The high-order filter is sized to minimizethe phase lag in the low frequency region (below 10 rad/sec) whileproviding more attenuation of the main and tail rotor resonant peakamplitudes than the baseline filters. Additionally, the anti-resonantnotch frequencies of the high-order filter are sized to account for theeffective torsional frequency shifts between the single and twin coupledengine operation.

The high-order filter of the subject invention can be configured asthree second order (notch) filters cascaded in series with two firstorder lead/lag filters. Alternatively, the high-order torsional filtercan be configured as a sixth order lead/lag filter in series with asecond order lead/lag filter.

Yaw anticipation in an engine control system is designed to accommodatethe change in rotor load resulting immediately from changes in tailrotor pitch. Similarly, lateral cyclic anticipation in an engine controlsystem is designed to accommodate the change in rotor load resultingfrom right/left lateral rolls. The increased bandwidth of the outerspeed loop 12 of control system 10 eliminates the need for yaw andlateral cyclic open loop load anticipation. It has been determinedhowever, that open loop anticipators are necessary for rotor decayanticipation and collective pitch anticipation. Accordingly, the controlsystem of the subject invention does incorporate appropriate rotor loadanticipators.

Optional Torsional Damping Loop

Referring once again to FIG. 1, the engine control system 10 of thesubject invention includes an optional active inner torsional dampingloop 14. Damping loop 14 receives a signal indicative of the measuredpower turbine shaft torque (QS) from a sensor operatively associatedwith the power turbines and the rotor drive train. The measured signalinput into a summing junction 50 which also receives an estimated QSwhich is a function of gas generator speed (NG). The difference betweenthe actual QS and the estimated QS, otherwise referred to as the ΔQSfrom the summing junction, is input into a high pass filter 52. Thishigh pass filter effectively nulls out any steady state errors in thesignal, and allows only the torsional frequencies to pass therethrough.

The filtered signal is then sent to a Kalman state estimator 54 which isprogrammed to estimate twelve different dynamic variables associatedwith the power turbines and rotor drive train using the measured QS toestimate all of the other states or dynamic variables for feedbackcontrol. The twelve states estimated by the Kalman estimator are definedas follows:

-   -   ₁=Main Rotor Gearbox Speed (ΔNGB)    -   ₂=Main Rotor Speed (ΔNR)    -   ₃=Effective Lag Hinge Damper Velocity (ΔNLHD)    -   ₄=Main Rotor Shaft Torque (ΔQRS)    -   ₅=Tail Rotor Speed (ΔNT)    -   ₆=Gas Generator High Pressure Turbine Speed (ΔNH)    -   ₇=Engine Burn Flow (ΔWfburn)    -   ₈=Power Turbine Speed (ΔNF)    -   ₉=NDOT Control Output (ΔWF/P3)    -   ₁₀=HMU Output Fuel Flow (ΔWF)    -   ₁₁=Tail Rotor Shaft Torque (ΔQTS)    -   ₁₂=Power Turbine Shaft Torque (ΔQS)

The twelve estimated states are input into a Linear Quadratic Regulator(LQR) 56 where they undergo gain reduction/amplification and summation,as described in more detail hereinbelow with respect to FIG. 5. Theresulting control signal is indicative of ΔNDOT* and is passed tosumming junction 20 where it is added to the power turbine governoroutput signal and the signal from the rotor load anticipators. As notedabove, the active damping loop 14 is optional. Accordingly, the innerloop 14 includes a switching gate 58 that follows the LQR summingjunction, that may be selectively actuated to activate the damping loop.

The functional operation of the LQR and Kalman estimator is illustratedin FIG. 5. As shown, the measured raw shaft torque and the estimatedshaft torque are summed to produce a ΔQS. The ΔQS is input into a highpass K_(T)QS filter which effectively nulls out steady state error inthe shaft torque signal and generates a filtered ΔQS which is summedwith the unfiltered ΔQS to produce a shaft torque signal indicative ofthe torsional frequencies of the rotor drive system.

The resultant signal is applied to a (12×1) Kalman gain vectorcontaining the twelve estimated dynamic state variables discussed above.The resulting values are input into a Kalman summing junction whichreceives three additional feedback signals indicative of various dynamicproperties of the control system architecture (A, B and C). The summingjunction produces an

DOT vector that is input into a (12×1) integrator to generate a signalindicative of

.

The three additional feedback signals sent to the Kalman summingjunction follow from the integrated

. The first feedback loop includes a (12×1) output vector C and a (1×12)Kalman gain vector. The second feedback loop includes a (1×12) K or LQRgain vector, a gain reducer and a (12×1) input vector B. The thirdfeedback loop includes a (12×12) dynamic property vector A. The signalsfrom the three feedback loops and summed at the Kalman summing junctionwith the primary Kalman gain vector. The output signal from the

integrator is summed and reduced as a function of gas generator speed(NG), resulting in the aforementioned ΔNDOT*.

The design of the Linear Quadratic Regulator (LQR) is based upon thefollowing methodology. For the linearized system defined by thefollowing equations: $\overset{.}{x} = {{Ax} + {B\;\mu}}$ y = Cx

-   -   where:        -   χ=Plant (engine) States        -   μ=NDOT Demand        -   y=Output of Interest, i.e., QS

The optimal regulator problem is to find state feedback control μ=−Kχsuch that the cost function:$J = {\int_{0}^{\infty}{\frac{1}{2}\left( {{x^{T}{Qx}} + {\mu^{T}R\;\mu}} \right)\ {\mathbb{d}t}}}$is minimized where Q and R are state and control weighting functions,respectively. In this problem, Q must be chosen so as to obtain acontrol that provides combustive damping. Thus, Q must be chosen so asthe total torsional energy in the rotor drive train, which isrepresented by the following equation:$E = {{\frac{1}{2}J_{R}N_{R}^{2}} + {\frac{1}{2}K_{R}\vartheta_{R}^{2}} + {\frac{1}{2}J_{T}N_{T}^{2}} + {\frac{1}{2}K_{T}\vartheta_{T}^{2}}}$where θR and θT are the main and tail rotor twist angles, respectively,N_(R) and N_(T) are the main and tail rotor speeds, respectively; K_(R)and K_(T) are the main and tail rotor stiffness factors, and J_(R) andJ_(T) are the main and tail rotor moments of inertia.

Since${\vartheta_{R} = \frac{Q_{R}}{K_{R}}},{{\vartheta_{T} = \frac{Q_{T}}{K_{T}}};}$It follows that the energy functional equation is:$E = {\frac{1}{2}\left( {{J_{R}N_{R}^{2}} + \frac{Q_{R}^{2}}{K_{R}} + {J_{T}N_{T}^{2}} + \frac{Q_{T}^{2}}{K_{T}}} \right)}$and it is expressed in terms of the rotor system dynamic states. Bysolving the LQR problem using conventional techniques and using thefunction total energy in the rotor system, a control is obtained thatdrives the rotor system energy to zero in the shortest possible time,resulting in combustive damping. Furthermore, an appropriate choice ofcontrol weight R will yield a controller that is a trade-off betweenperformance (i.e., the amount of damping) and stability margin.

The Kalman estimator is designed as an adjunct problem to the LQRproblem whereby state weight Q is selected to be the identity matrix sothat the estimated states are driven to be equal to actual values, andcontrol weight R is chosen to give acceptable bandwidth (e.g., 70rad/sec) and to attenuate higher frequencies.

Although the control system of the subject invention has been describedwith respect to preferred embodiments, those skilled in the art willreadily appreciate that changes and modifications may be made theretowithout departing from the spirit and scope of the present invention asdefined by the appended claims.

1. A power turbine speed control system for a helicopter comprising: a)means for generating a power turbine speed signal based upon a demandedspeed; b) means for filtering the power turbine speed signal byeffectuating a rapid attenuation of main and tail rotor torsionalfrequencies in the power turbine speed signal without compromising phaseat low frequencies; c) a governor for providing isochronous powerturbine speed and rotor speed control based upon the filtered powerturbine speed signal; and d) damping means for actively damping main andtail rotor torsional frequencies.
 2. A power turbine speed controlsystem as recited in claim 1, wherein the damping ineans includes meansfor estimating a plurality of engine states based upon a single measuredengine state.
 3. A power turbine speed control system as recited inclaim 2, wherein the single measured engine state is power turbine shafttorque.
 4. A power turbine speed control system as recited in claim 1,wherein the damping means includes a linear quadratic regulator thatprovides combustive damping.
 5. A power turbine speed control system asrecited in claim 1, further comprising means for selectively activatingthe damping means.
 6. A turbine speed control system as recited in claim1, wherein the damping means is tuned to provide attenuation of resonantfrequencies so as not to influence low frequency response of the system.7. A power turbine speed control system as recited in claim 6, whereinthe damping means includes a high pass filter.